Process for operating a dual-mode combustor

ABSTRACT

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

ORIGIN OF THE INVENTION

The invention described herein was made by employees of the UnitedStates Government and may be manufactured and used by or for theGovernment for Government purposes without the payment of any royaltiesthereon or therefor.

CROSS REFERENCE TO RELATED APPLICATION

This patent application is a Divisional and claims priority to U.S.patent application Ser. No. 12/894,346 filed Sep. 30, 2010, now U.S.Pat. No. 8,484,980.

FIELD OF THE INVENTION

The invention is in the field of dual-mode combustors for use as aramjet and a scramjet.

BACKGROUND OF THE INVENTION

Combined-cycle propulsion is considered when the high efficiency ofair-breathing propulsion is desired over a broad Mach number range.Air-breathing access to space is one such application of currentinterest to NASA. The dual-mode scramjet is central to mostcombined-cycle schemes. Turbine-based combined-cycle (TBCC) systems usea turbine engine for low speed acceleration, and operate to a maximumflight Mach number in scramjet mode dictated by system considerations.TBCC systems are normally assumed to take-off horizontally, and use asecond, rocket-powered stage to achieve orbit. Rocket-basedcombined-cycle (RBCC) schemes use chemical rocket propulsion for lowspeed acceleration. The high thrust-to-weight ratio of the rocketcomponent allows for its integration within the air-breathing duct. RBCCsystems are normally assumed to be launched vertically, and can operatefrom lift-off to orbit. Turbine-engines reach temperature and thrustlimitations as Mach number increases. Rocket thrusters provide a highratio of thrust-to-weight at any speed, but are very inefficient fromthe standpoint of specific impulse. In either case, it is advantageousto extend dual-mode scramjet operation to as low a Mach number aspossible.

Supersonic combustion has long been recognized as a solution to problemsassociated with the severe stagnation conditions within a ramjet engineat high flight Mach number. Diffuser momentum loss, dissociation,non-equilibrium expansion losses, and structural loading are allrelieved by transition to a supersonic combustion process. In general,the cross-sectional area of the supersonic combustor increases in thedownstream direction to avoid thermal choking and excessive pressuregradients. The subsequent nozzle expansion process requires a moredramatic increase in cross-sectional area and is usually integrated withthe vehicle aft end to provide the maximum possible area ratio.

In order to extend the operable flight Mach number range of the scramjetengine downward, toward the upper limit for turbojets or to limit rocketoperation to as low a ΔV (speed range) as possible, “dual-mode”operation was introduced by Curran, et al. in U.S. Pat. No. 3,667,233.U.S. Pat. No. 3,667,233 is incorporated herein by reference hereto.

FIG. 1 is a prior art drawing from Curran et al., U.S. Pat. No.3,667,233, and, in particular, is a schematic diagram partially in blockform of a dual mode combustion chamber according to the invention.

FIG. 2 is a prior art drawing from Curran et al., U.S. Pat. No.3,667,233, and, in particular, is a schematic cross section of thedevice of FIG. 1 showing one possible arrangement of the fuel injectors.

FIG. 3 is a prior art drawing from Curran et al., U.S. Pat. No.3,667,233, and, in particular, is a schematic diagram partially in blockform showing an annular configuration for the combustion chamber of FIG.1.

FIG. 4 is a prior art drawing from Curran et al., U.S. Pat. No.3,667,233, and, in particular, is a schematic end view of the device ofFIG. 3 from the exhaust end.

FIG. 5 is a prior art drawing from Curran et al., U.S. Pat. No.3,667,233, and, in particular, is a schematic diagram partially in blockform of a modified fuel supply system for the device of FIG. 1.

Conceptually, a thermally-choked combustion process is established inthe aft regions of the scramjet flowpath where the cross-sectional areasare greatest. As depicted in FIG. 7, the diverging scramjet duct acts asa subsonic diffuser, and the thermal throat is positioned so as toproduce the required back-pressure.

FIGS. 6 and 7 are another illustration of the structure and process ofthe prior art Curran et al., U.S. Pat. No. 3,667,233.

Curran et al., U.S. Pat. No. 3,667,233, states at col. 1, lns. 29 etseq. that:

“A combustor with a fixed geometry has one parallel combustion sectionwith a substantially uniform cross section along its length. Fuel isinjected into this section and the flame is stabilized on recessedflameholders. As the fuel burns it causes choked flow in this sectionwhich sends a shock wave upstream to convert the normal supersonic flowthrough the combustor to subsonic flow. For transition from subsonicmode to the supersonic mode, fuel is injected into a diverging sectionupstream of the parallel section which causes the shock to movedownstream until it is ejected from the engine. In the final transitionto supersonic mode, fuel is supplied only to the upstream injectors.”

Further, Curran et al., U.S. Pat. No. 3,667,233, states at col. 2, lns.4 et seq. that:

“At these speeds fuel is supplied to nozzles 36. Burning of the fuel inthe uniform cross section combustion chamber 24 causes choked flow whichsends a shock wave upstream of the flow to convert the supersonic flowto subsonic flow within the combustion chamber. As the speed of theaircraft increases to a speed between Mach 4 and Mach 5, fuel control 30starts a flow of fuel to nozzles 32 as the fuel control 34 graduallydecreases the fuel flow to nozzles 36. This causes the shock wave togradually recede as fuel to nozzles 32 is increased and fuel flow isdecreased to nozzles 36. At a speed of about Mach 8 fuel to nozzles 36is further reduced and supersonic combustion now occurs throughout thedivergent and parallel ducts. The expansion of the heated gases inexpansion section 22 permits higher Mach speeds to be attained.”

The cross-sectional area of the thermal throat must increase as flightMach number decreases, unless fuel-to-air ratio is reduced. For a givenduct, this effect determines the minimum flight Mach number fordual-mode operation. At Mach 2.5, the required thermal throat areaapproaches that of the inlet capture area. The primary technicalchallenges in practical application of the dual-mode scramjet scheme ofCurran et al., U.S. Pat. No. 3,667,233, are modulation of the thermalthroat location, modulation of fuel distribution, ignition, andflame-holding in the large cross-section. Any in-stream devices must beretractable or expendable so as not to inhibit supersonic combustionoperation.

Curran et al., U.S. Pat. No. 3,667,233, controls fuel flow to modulatethe position of the thermal throat at low flight Mach numbers and then,subsequently, to transition to supersonic ramjet operation. If Currandoesn't modulate the position of the choked flow correctly, the shockwave moves further upstream into the inlet passage 21 of Curran andun-start of the engine may occur.

FIG. 6 shows a cross-sectional view 600 of a prior art (Curran et al.)scramjet engine operating in the scramjet mode. Processes that governscramjet efficiency are inlet momentum losses, Rayleigh losses due toheat addition, heat loss to the combustor walls, skin friction, andnon-equilibrium expansion. Other factors that must be considered includeseparation of boundary layers due to adverse pressure gradients, intenselocal heating at re-attachment points and shock impingements, and fuelstaging or variable geometry to accommodate the variation of combustionarea ratio with free stream stagnation enthalpy.

Referring to FIG. 6, fuel injection nozzle 601, inlet contractionsection 602, diverging supersonic combustion section 603, and exitnozzle 604 are illustrated. As stated above, in the scramjet mode, thisengine is fed with fuel injector 601. Reference numeral 608 illustratesand internal wall of the engine. Reference numeral 606 signifiesincoming air being compressed. Reference numeral 605 represents thefuel-air mixture being combusted. And, reference numeral 607 signifiesexpanded gas/combustion products being expelled from the engine.

FIG. 7 is the cross-sectional view 700 of FIG. 6 (Curran et al. priorart engine) in the ramjet mode illustrating choked flow 702 and a shockwaves 701. Fuel injectors 703, 704 are illustrated and are operable inthe ramjet mode. Curran et al. must delicately control the insertion offuel. First, fuel is inserted with injectors 703, 704 and then fuel isinserted using injector 601 to prevent the shock wave from beingexpelled leftwardly into the inlet contraction section 602 which mayresult in un-start of the engine. Reference numeral 606A indicatesincoming compressed air and reference numeral 607A represents combustionproducts expelled from the engine.

SUMMARY OF THE INVENTION

The supersonic free jet mode of a new combined-cycle combustor isdisclosed herein at various scramjet flight Mach numbers including 5, 8,and 12. The dual-mode combustor has the ability to operate in ramjetmode to lower flight Mach numbers than current dual-mode scramjets,thereby bridging the gap between turbine or rocket-based low speedpropulsion and scramjets.

One important feature of the invention is the use of an unconfined freejet for supersonic combustion operation at high flight mach numbers. Thefree jet traverses a larger combustion chamber that is used for subsoniccombustion operation at lower flight Mach numbers. The free jet expandsat constant pressure due to combustion and rejoins the nozzle throatcontour. Recirculating flow in the combustion chamber equilibrates to apressure slightly lower than that of the free jet causingunder-expansion features to appear in the free-jet. The free jet joinsthe nozzle throat contour with little interaction and expands throughthe nozzle expansion section.

At scramjet flight Mach numbers from 5 to 12, the supersonic free jettraverses the combustion chamber and rejoins the nozzle contour at thecombustor exit. Periodic wave structure occurs in the free-jet and isinitiated by an entry interaction caused by pressure mismatch and rapidmixing and combustion at the combustion chamber entrance and upstream inthe inlet section. The periodic nature of the free jet also led to anexit interaction determined by the phase of the wave structure withrespect to the throat location. The effect of reducing nozzle throatarea was to increase the combustion chamber pressure, and reduce theperiod of the wave structure, but not its amplitude. A viscous loss dueto momentum transfer to the recirculation zone is also apparent in eachcase.

Calculated heat loads were commensurate with previous estimates for airbreathing systems. Peak heat flux occurred upstream of the throat at animpingement point separating the free-jet from recirculation zone. For agiven wall temperature, heat load depends on the recirculation zonetemperature and volume, the severity of the exit interaction, and thefuel injection scheme.

The new combustor is disclosed for use over a wide range of flight Machnumbers, operating in both subsonic and supersonic combustion modes. Itoperates as a conventional ramjet at low speed, eliminating theaforementioned issues with dual-mode operation. Transition to supersoniccombustion in a free jet mode occurs at the appropriate flight conditionupon the rapid opening of the nozzle throat.

A supersonic combustion ramjet engine is disclosed and claimed. Theterms supersonic combustion ramjet engine, supersonic combustion ramjetand dual-mode combustor are used interchangeably herein. The supersoniccombustion ramjet engine is operable in a ramjet mode and a scramjetmode. The ramjet mode extends from about flight Mach number 2.5 up toabout flight Mach number 6. The scramjet mode extends from about flightMach number 5 up to about flight Mach number 12. An inlet passagewayreceives compressed combustion air from a supersonic diffuser. The inletpassageway includes a fuel injector. A subsonic diffuser and acombustion chamber follow the inlet passageway. The subsonic diffuser(sometimes referred to herein as the diffusion section) includes aninner periphery. A radial step is interposed between and links the inletpassageway and the diffusion section.

The inlet passageway is in communication with the diffusion section andthe diffusion section is in communication with the combustion chamber. Aramjet-mode flame holder array is located between the subsonic diffuserand the combustion chamber. The flame holder array includes a centralcircular aperture therethrough. The flame holders are affixed to theinner periphery of the combustion chamber.

The engine also includes a contraction section, a variable nozzle throatand an expansion section. The combustion chamber is in communicationwith the nozzle contraction section and the nozzle contraction sectionis in communication with the variable nozzle throat. And, the variablenozzle throat is in communication with the expansion section.

A nozzle positioner drives and moves the arc section forming thevariable nozzle throat to a desired diametrical opening according to analgorithm which is a function of flight Mach number and combustor mode.The algorithm has a discontinuity at a given flight mach numbertransitioning from the ramjet mode to the scramjet mode forming afree-jet from the inlet section, through the subsonic combustion chamberand reattaching at the variable nozzle throat. The ramjet mode includessubsonic operation from about flight Mach number 2.5 up to about flightMach number 5.0 to 6.0. The scramjet mode includes supersonic operationfrom about flight Mach number 5.0 to 6.0 up to about flight Mach number12.0 and greater. The nozzle positioner divergingly adjusts the nozzlethroat to a relatively larger diameter between about flight Mach number5.0 to 6.0 transitioning from the ramjet mode to the scramjet modeforming a free jet extending from the inlet section at the location ofthe radial step to the variable nozzle throat. The free-jet does notengage the subsonic diffuser. Nor does the free jet engage thecombustion chamber. The free-jet rejoins the variable nozzle throat.

A supersonic diffuser is used to compress combustion air into acombustion air passageway. Fuel is injected from the combustion airpassageway into the combustion air in the combustion air passagewaycreating a stoichiometric fuel-air mixture. In the scramjet mode, thestoichiometric fuel-air mixture is fed from the combustion airpassageway into a free jet that traverses the subsonic diffuser.Operation in the scramjet mode is premised on previous operation andignition in the ramjet mode using flame holders in the subsonicdiffuser.

In ramjet mode, fuel is in injected from the combustion air passageway.In the ramjet mode, the fuel-air mixture is combusted in the combustionchamber. The combusted fuel-air mixture is evacuated from the combustionchamber and into the variable area nozzle throat. The variable nozzlethroat is modulated and positioned according to an algorithm creatingand controlling the position of a terminal shock in the subsonicdiffuser. The algorithm is a function of flight Mach number.

The step of discontinuing operation of the flame holder, and the step ofmodulating and positioning a variable nozzle throat according to analgorithm, includes transitioning, using the algorithm, the dual-modecombustor from a ramjet mode to a scramjet mode by rapidly opening thevariable nozzle throat at a specified flight Mach number. The algorithmincludes a discontinuity where there are two values for a specifiedflight mach number and it is this discontinuity, and the action basedupon it, which shifts the dual-mode combustor from the ramjet mode tothe scramjet mode. Shifting from the scramjet mode to the ramjet mode isalso possible.

The algorithm includes the variable nozzle position as a ratio A/Ac ofthe actual nozzle throat area, A, to the inlet capture area, Ac, of thesupersonic diffuser. The nozzle position varies from a ratio of about0.8=A/Ac at about flight Mach number 2.5 in the ramjet mode to a ratioof about 0.18=A/Ac at about flight Mach number 5.0 in the ramjet mode.The nozzle position varies rapidly from a ratio of about 0.18=A/Ac atabout flight Mach number 5.0 in the ramjet mode to a ratio of about0.41=A/Ac at about flight Mach number 5.0 transitioning to the scramjetmode. Thereafter, the nozzle position varies from about 0.41=A/Ac atflight Mach number 5.0 in the scramjet mode to a ratio of about0.15=A/Ac at about flight Mach number 12 in the scramjet mode.

In the scramjet mode, the fuel-air mixture and the combustion productsare separated into a free-jet beginning at the exit of the combustionair passageway/radial step and extends to the variable nozzle throat.The free-jet does not engage the subsonic diffuser, the combustionchamber or the contraction section.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and,in particular, is a schematic diagram partially in block form of a dualmode combustion chamber according to the invention.

FIG. 2 is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and,in particular, is a schematic cross section of the device of FIG. 1showing one possible arrangement of the fuel inlet jets.

FIG. 3 is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and,in particular, is a schematic diagram partially in block form showing anannular configuration for the combustion chamber of FIG. 1.

FIG. 4 is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and,in particular, is a schematic end view of the device of FIG. 3 from theexhaust end.

FIG. 5 is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and,in particular, is a schematic diagram partially in block form of amodified fuel supply system for the device of FIG. 1.

FIG. 6 is another cross-sectional view of the prior art Curran deviceoperating in scramjet mode.

FIG. 7 is the cross-sectional view of FIG. 6 in the ramjet modeillustrating choked flow and a shock wave.

FIG. 8 is a perspective view of the dual-mode combustor in the ramjetmode.

FIG. 8A is a cross-sectional schematic view of the dual-mode combustorof FIG. 8 in the ramjet mode.

FIG. 8B is a quarter-sectional schematic view of the dual-mode combustorof FIG. 8 in the ramjet mode.

FIG. 8C is an enlargement of a portion of FIG. 8B illustrating,diagrammatically, the step between the inlet cylinder and the subsonicdiffuser.

FIG. 9 is a perspective view of the dual-mode combustor in the scramjetmode.

FIG. 9A is a cross-sectional schematic view of the dual-mode combustorof FIG. 9 in the scramjet mode illustrating the free jet extending fromthe inlet cylinder to the variable nozzle throat.

FIG. 9B is a quarter-sectional schematic view of the dual-mode combustorof FIG. 9 in the scramjet mode.

FIG. 9C is a sectioned view of the dual-mode combustor of FIG. 9illustrating the flame-holder having a central aperture therein for thepassage of the free-jet.

FIG. 10 is another example of the dual-mode combustor employingdifferent geometry.

FIG. 11 is a quarter-sectional diagrammatic view of the dual modecombustor in the scramjet mode for flight mach number 8 illustrating astep, a hinged diffuser section, a fixed combustion chamber section, ahinged contraction section, a hinged nozzle throat section (circular arcsection) and a hinged expansion section.

FIG. 11A is a table of values relating to FIG. 11.

FIG. 11B is a schematic view of an example of a receiving joint formingthe nozzle throat.

FIG. 12 is a plot of the prior art thermal throat of Curran, thegeometric/nozzle throat of the dual-mode combustor in ramjet mode and inscramjet mode, and the inlet diameter of the dual-mode combustor as afunction of the inlet capture area.

FIG. 12A is a table of inlet contraction ratios as a function of theinlet capture area for a range of flight Mach numbers.

FIG. 12 B is a control system for positioning the variable (geometric)nozzle throat.

FIG. 13A is a generalized quarter-sectional diagrammatic view of theflight Mach number 2.5 ramjet.

FIG. 13B is a generalized quarter-sectional diagrammatic view of theflight Mach number 3.0 ramjet.

FIG. 13C is a generalized quarter-sectional diagrammatic view of theflight Mach number 4.0 ramjet.

FIG. 14A is a generalized quarter-sectional diagrammatic view of theflight Mach number 5.0 ramjet.

FIG. 14B is a generalized quarter-sectional diagrammatic view of theflight Mach number 5.0 scramjet.

FIG. 15A is a generalized quarter-sectional diagrammatic view of theflight Mach number 6.0 ramjet.

FIG. 15B is a generalized quarter-sectional diagrammatic view of theflight Mach number 6.0 scramjet.

FIG. 16A is a generalized quarter-sectional diagrammatic view of theflight Mach number 8.0 scramjet.

FIG. 16B is a generalized quarter-sectional diagrammatic view of theflight Mach number 10.0 scramjet.

FIG. 16C is a generalized quarter-sectional diagrammatic view of theflight Mach number 12.0 scramjet.

FIG. 17A is an illustration of the pressure contours within the enginefor the flight Mach number 5.0 scramjet.

FIG. 17B is an illustration of the pressure contours within the enginefor the flight Mach number 8.0 scramjet.

FIG. 17C is an illustration of the pressure contours within the enginefor the flight Mach number 12.0 scramjet.

FIG. 18A is an illustration of the Mach number contours within theengine for the flight Mach number 5.0 scramjet.

FIG. 18B is an illustration of the Mach number contours within theengine for the flight Mach number 8.0 scramjet.

FIG. 18C is an illustration of the Mach number contours within theengine for the flight Mach number 12.0 scramjet.

FIG. 19A is an illustration of the static temperature contours withinthe engine for the flight Mach number 5.0 scramjet.

FIG. 19B is an illustration of the static temperature contours withinthe engine for the flight Mach number 8.0 scramjet.

FIG. 19C is an illustration of the static temperature contours withinthe engine for the flight Mach number 12.0 scramjet.

FIG. 20 illustrates ideal net thrust per unit airflow against flightMach numbers for a conventional ramjet, thermally-choked ramjet and ascramjet.

FIG. 21 illustrates the mass-averaged static pressure distributions withthe pressure at the nozzle throat station (supersonic combustor exit)denoted by symbols.

FIG. 21A illustrates the mass-averaged axial velocity distributions forvarious flight conditions.

FIG. 21B illustrates the mass-averaged temperature distributions forflight Mach numbers 5, 8 and 12.

FIG. 22A illustrates the static pressure ratio for scramjet mode flightMach number 8 with the variable nozzle throat positioned at 110% of thedesign operating point.

FIG. 22B illustrates the static pressure ratio for scramjet mode flightMach number 8 with the variable nozzle throat positioned at 100% of thedesign operating point.

FIG. 22C illustrates the static pressure ratio for scramjet mode flightMach number 8 with the variable nozzle throat positioned at 90% of thedesign operating point.

FIG. 22D illustrates the static pressure ratio for scramjet mode flightMach number 8 with the variable nozzle throat positioned at 80% of thedesign operating point.

FIG. 23 illustrates the effect of nozzle throat area variation forscramjet mode flight Mach number 8 on the rate of ethylene fueldepletion.

FIG. 23A illustrates the ethylene mass fraction for flight Mach numbers5, 8 and 12 versus axial position.

FIG. 24 illustrates the effect of nozzle throat area variation onmass-averaged static pressure distribution for scramjet mode flight Machnumber 8.

FIG. 25 illustrates the ideal net thrust per unit of airflow plottedagainst combustor exit pressure ratio and nozzle throat area variationfor scramjet mode flight Mach number 8.

DESCRIPTION OF THE INVENTION

In the design of the dual-mode combustor all processes were assumedadiabatic. Air capture, inlet contraction ratio, and total pressurerecovery were specified as a function of flight Mach number illustratedin FIG. 12A. These characteristics are representative of a single-cone,axi-symmetric inlet design with forebody pre-compression. Air wasassumed to be a mixture of nitrogen and oxygen at 78.85% and 21.15% byvolume, respectively.

In the analysis of all ramjet cases, ethylene fuel entered at sonicvelocity, normal to the propulsion axis at 5180 R. The energy requiredto raise the ethylene fuel to this condition was ignored. Constant-areacombustion in a cross-sectional area equal to 83.3% of the inlet capturearea was assumed. This area was chosen to allow operation at a minimumflight Mach number of 2.5 without thermal choking. For comparison,calculations were also done assuming a thermally-choked combustionprocess. For these cases, the diffuser exit Mach number was set toresult in a combustion area ratio of 1.5.

The AIAA (American Institute of Aeronautics and Astronautics), paperentitled Supersonic Free-Jet Combustion in a Ramjet Burner, by CharlesJ. Trefny and Vance F. Dippold III, NASA Glenn Research Center,Cleveland, Ohio, 44135 was published and presented on Jul. 26, 2010 isincorporated herein by reference hereto.

The dual-mode combustor is illustrated in FIGS. 9, 9A and 9B in scramjetmode wherein supersonic combustion in an unconfined free-jet 943traverses a larger subsonic combustion chamber 805, a contractionsection 806, and a variable nozzle throat 807. FIG. 9 is a perspectiveview 900 of the dual-mode combustor in the scramjet mode. FIG. 9A is across-sectional schematic view 900A of the dual-mode combustor 899 ofFIG. 9 in the scramjet mode illustrating the free jet 943 extending fromthe inlet cylinder 802 to the variable nozzle throat 807 which yields anozzle throat diameter D1. Nozzle throat diameter D1 as illustrated inFIG. 9A is larger than nozzle throat diameter D illustrated in FIG. 8A.The nozzle throat area is dictated by the curves illustrated in FIG. 12for both the ramjet and the scramjet. The examples of nozzle position807 given here for the ramjet and the scramjet are the portions of thecurves 1202, 1205 where FIG. 8A ramjet mode uses a smaller nozzle throatarea than FIG. 9A (scramjet mode).

FIG. 9B is a quarter-sectional schematic view 900B of the dual-modecombustor of FIG. 9 in the scramjet mode. In the scramjet mode,reference numeral 845A signifies supersonic combustion and referencenumeral 847A represents expansion. FIG. 9C is a sectioned view 900C ofthe dual-mode combustor of FIG. 9C illustrating the flame-holder 810having a central aperture 850 therein for the passage of the free jet943 there through.

During scramjet mode of operation, the propulsive stream 943 is not incontact with the combustor walls 805, and equilibrates 943A to thecombustion chamber pressure 944. Boundary 943A represents the interfaceof the free-jet/propulsive stream 943 with the recirculationzone/combustion chamber pressure 944. Thermodynamic efficiency issimilar to that of a traditional scramjet under the assumption ofconstant-pressure combustion. Qualitatively, a number of possiblebenefits exist. Fuel staging is eliminated since the cross-sectionalarea distribution required for supersonic combustion is accommodatedaerodynamically without regard for wall pressure gradients andboundary-layer separation because the free-jet does not touch the wallsof the diffuser and the combustion chamber. Variable exit diameter D1must be set to the proper size for a given flight Mach number. The axialdistance available for supersonic mixing and combustion includes thesubsonic diffuser 804, combustion chamber 805 and nozzle contractionsections 806 required for ramjet operation. Heat loads, especiallylocalized effects of shock-boundary-layer interactions, are reduced.Reference numeral 880 signifies incoming air being compressed andreference numeral 881 signifies exiting combustion gases.

FIG. 8 is a perspective view 800 of the dual-mode combustor 899 in theramjet mode. FIG. 8 illustrates the frusto-conical inlet contractionsection 801, the cylindrical inlet passageway 802, the diffuser section804, the combustion chamber 805, the contraction section 806 and thevariable diameter nozzle throat 807. Reference numeral 807 signifies thevariable nozzle throat at the joining point of the contraction section806 and the expansion section 808 in the ramjet mode. In the scramjetmode, reference numeral 807 also signifies the variable nozzle throat atthe joining point of the contraction section 806 and the expansionsection 808.

FIG. 8A is a cross-sectional schematic view 800A of the dual-modecombustor 899 of FIG. 8 in the ramjet mode. FIG. 8A illustratessubstantial differences in construction when compared to Curran U.S.Pat. No. 3,667,233. First, the flame holders 810 are arranged so as tonot obstruct the free-jet as illustrated in FIG. 9A. The flame holders810 have a central, circular aperture 850 therein. Reference numeral810A signifies the flame holders in operation. Reference numeral 830represents a terminal shock wave and its location as illustrateddiagrammatically in FIG. 8A is important. Location of the terminal shockwave 830 in the ramjet mode is important and is controlled by theposition of the nozzle throat 807 diameter D. Reference numeral 872signifies heat release within the combustor.

There is no thermal throat in the dual-mode combustor 899 because thevariable nozzle throat 807 is positioned so as to control the terminalshock wave 830.

FIG. 12 is a plot 1200 of the prior art thermal throat of Curran 1201,the geometric/nozzle throat 1202 of the dual-mode combustor 899 inramjet mode, the geometric/nozzle throat 1205 in scramjet mode, and theinlet throat 1203, 1203A of the dual-mode combustor 899 as a ratio ofA/A capture area. FIG. 12A is a table 1200A of inlet contraction ratios1231 as a ratio ((A/A capture area) 1231) for a range of flight machnumbers and combustion processes 1230. FIG. 12 B is a control system1200B for positioning the variable (geometric) nozzle throat 807. FIG.12 indicates a discontinuity or jump 1204 between the ramjet mode plot1202 and the scramjet mode plot 1205.

A nozzle positioner 1212 drives and moves the arc section 1125 formingthe nozzle throat 1108, to a desired diametrical opening according to analgorithm (FIG. 12 curves, 1202, 1205) which is a function of flightMach number and combustor mode (ramjet or scramjet). The algorithm has adiscontinuity at a given flight mach number, in this example, flightMach number 5.0, transitioning from the ramjet mode to the scramjet modeforming a free-jet 943 from the inlet section 802, through the subsonicdiffuser 804, through the combustion chamber 805, through thecontraction section, and rejoins the nozzle throat 807 (diameter D1).The ramjet mode includes subsonic operation from about flight Machnumber 2.5 up to about flight Mach number 5.0 to 6.0 and thecross-sectional area of the nozzle throat 807 divided by the inletcapture area, A inlet capture area, should follow curve 1202.

The scramjet mode includes supersonic operation from about flight Machnumber 5.0 to 6.0 up to about flight Mach number 12.0 and greater. Thenozzle positioner divergingly adjusts the nozzle throat diameter (nozzlearea) rapidly to a relatively larger diameter between about flight Machnumber 5.0 to 6.0 rapidly transitioning from the ramjet mode to thescramjet mode forming a free-jet 943 extending from the inlet section802 at the location of the radial step 812, 812A to the nozzle throat807. The free-jet does not engage the subsonic diffuser 804. Nor doesthe free-jet 943 engage the combustion chamber 805. The free-jet 943rejoins the nozzle throat 807 as illustrated in FIG. 9A.

Referring to FIG. 8A and FIG. 12, reference numeral 1201 indicates thealgorithm for the position of the nozzle throat 807 (diameter D) as aratio of the inlet capture area (area=A inlet capture area).Specifically, the nozzle throat area must be positioned on the line 1202for ramjet mode operation for flight numbers between 2.5 to 5.0.Further, the nozzle throat 807 (diameter D1) in the scramjet mode mustbe positioned on the line 1205 for the scramjet mode operation forflight numbers between 5.0 and 12.0. Reference numeral 1204 representsthe transition between the ramjet mode (pursuant to curve or algorithm1202) and the scramjet mode (pursuant to curve or algorithm 1205).Operation between the modes is switched back and forth between thecurves 1202, 1205.

Referring to FIG. 8A and FIG. 12, the location of the shock wave 830 isimportant. If the nozzle throat area ratio is positioned below the line1201 in FIG. 12, the engine will un-start as the shock wave movesleftwardly and is expelled from the engine in order to spill air aroundand past the inlet capture area. Similarly, if the nozzle throat arearatio is positioned above the line 1202 in FIG. 12, the engine mayprematurely transition to the scramjet mode if the flight Mach number issufficiently high. Transition to the scramjet mode is accomplished byrapidly changing the nozzle throat ratio (A/A inlet capture area) fromcurve 1202 to curve 1205 in combination with radially oriented step 1203which causes the free jet to separate from the diffuser surface and thecombustion chamber. The flame holders 810 have no function.

FIGS. 12 and 12A also indicate that the diameter of the cylindricalinlet 802 changes as a function of ramjet mode (see curve 1203), andalso cylindrical inlet 802 changes as a function of scramjet mode (seecurve 1203A). FIG. 12A indicates that the inlet contraction ratio (Ainlet capture area/A inlet cylinder) increases as flight Mach numberincreases in the ramjet mode up to about flight Mach number 6.0.Further, FIG. 12A indicates that the inlet contraction ratio increasesas flight Mach number increases in the scramjet mode up to about flightMach number 12.0. FIG. 12 reference numerals 1203, 1203A represent theinverse of this data, in other words, the inlet throat diameter ratios(A inlet cylinder/A capture area) are the inverse of the previouslydefined contraction ratio.

As a general rule the nozzle throat 1202 and the inlet throat 1203decrease with increasing flight Mach number in ramjet mode. Similarly,as a general rule the geometric/nozzle throat 1205 and the inlet throat1203A decrease with increasing flight Mach number in scramjet mode.

Cycle analysis was performed over the flight Mach number range of 2.5 to12 at a dynamic pressure of 1500 psfa in order to establish the variablegeometry requirements for the inlet area and nozzle throat area. Forsupersonic combustion cases, a constant-pressure combustion process wasassumed with ethylene fuel entering at sonic velocity, parallel to thepropulsion axis at the diffuser exit static pressure and 10000 R.

FIG. 12 presents the variation of inlet and nozzle throat areas withflight Mach number for various operating modes. Of primary interest isthe large variation in nozzle throat area required in the low flightMach number range. The dual-mode ramjet's thermal throat area must varyby a factor of 4.5 from Mach 2.5 to 5. The required throat areavariation for the conventional ramjet is slightly less over the samerange. The thermally-choked cases require a larger throat area at agiven flight Mach number because of the greater total pressure lossassociated with the transonic combustion process. In the dual-modeengine, the axial location of combustion in a diverging flow path isvaried. The fuel distribution and flame-holding mechanisms used foraxial modulation of the heat release must not interfere withscramjet-mode operation. These are the fundamental issues associatedwith extension of the dual-mode to low Mach number flight. Also shown inFIG. 12 is the inlet throat area variation representative of thecontraction ratio. Finally, the combustor-exit area variation as aresult of constant-pressure supersonic combustion is shown in FIG. 12,and represents the free jet combustor nozzle throat area design values.

The area ratio due to combustion of the propulsive stream decreases withflight Mach number as the incoming energy increases. A factor of 2.5reduction in nozzle throat area is required between Mach 5 and 12. Forall modes of operation, the required variations in throat area shown area function of the inlet mass capture and pressure recoverycharacteristics assumed, and while representative for the purposesherein, could be reduced by integration, or other inlet design thatresults in greater spillage and higher recovery at the lower end of theflight Mach number range. Nozzle throat area variation requirementscould also be relieved by a reduction in fuel-air ratio at the lowerflight Mach numbers at the expense of net thrust. Obviously, limitingthe flight Mach number range would also diminish the variable geometryrequirements.

FIG. 12B is a control system 1200B for positioning the variable(geometric) nozzle throat 807. FIG. 12B illustrates desired 1206 ramjetmode (A nozzle/A inlet capture area) ratios switched into a controller1210 when in the ramjet mode. Similarly, FIG. 12B illustrates desired1208 scramjet mode (A nozzle/A inlet capture area) ratios switched intocontroller 1210 when in the scramjet mode. Controller 1210, based on anydifferences between desired and actual (A nozzle/A inlet capture area),outputs corrective action to the nozzle positioner 1212 which thenpositions 1214 the variable geometric nozzle throat. A nozzle positionersensor 1216 in combination with interconnecting lines 1215, 1217communicate the actual (A nozzle/A inlet capture area) signal tocontroller 1210 for comparison to the desired (A nozzle/A inlet capturearea) pursuant to curve or algorithm 1202, 1204 and 1205.

FIG. 8B is a quarter-sectional schematic view 800B of the dual-modecombustor 899 of FIG. 8 in the ramjet. mode. Supersonic compression 841occurs in the inlet contraction section 801 leading to the cylindricalinlet passageway 802. Arrow 842 indicates fuel injected perpendicularlyto the variable diameter inlet cylindrical passageway/section 802.Multimode fuel injector 842I injects fuel radially into passageway 802.Reference numeral 844 illustrates a region of subsonic diffusion andfuel mixing and reference numeral 845 illustrates a region of subsoniccombustion. Reference numeral 846 illustrates contraction to a chokedthroat 807 and reference numeral 847 illustrates expansion and exhaust.

FIG. 8C is an enlargement 800C of a portion of FIG. 8B illustrating,diagrammatically, the radial step 803 between the inlet 802 cylinder andthe subsonic diffuser 804. FIG. 8C also illustrates the fuel injector842I and the injection of fuel 842.

One of the important benefits of the dual-mode combustor 899, however,is that the combustion chamber 805 can be used for robust, subsoniccombustion at low flight Mach numbers. Operation as a subsoniccombustion ramjet (ramjet mode) is illustrated in FIGS. 8, 8A and 8B.Fuel injection can be accomplished with a single array of injectorsupstream in the inlet section 802. Ignition and flame-holding 810 can beaccomplished with an in-stream device as shown in FIGS. 8 and 9.

FIGS. 8, 8A and 8B illustrate the subsonic combustion ramjet mode. Atthe desired flight condition, transition to free jet mode is effected byincreasing the nozzle throat 807 area suddenly and inducing separationat the radial step 803 located at the diffuser inlet. The flame-holdingarray 810 does not extend across the subsonic diffuser 804. Inparticular, the flame-holding array includes an aperture 850 thereinwhich accommodates passage of the free jet therethrough in the scramjetmode. The subsonic diffuser section, sometimes referred to herein as thesubsonic diffuser 804, satisfies the requirements of operation as adiffuser in ramjet mode, and separated operation in free-jet mode.

In free-jet mode (scramjet mode) the propulsive stream re-joins thenozzle throat section, D1, with a minimum of disruption. The combustionchamber pressure equilibrates to near that of the diffuser exit, andwill depend on many factors such as the nozzle throat area, A, the rateof fuel entrainment, and the aerodynamics of the re-circulation region.Overall heat load to the combustion chamber walls depends on thetemperature in the recirculation region, and the competing effects oflow velocity and increased surface area.

FIG. 10 is a perspective of the dual-mode combustor 1000 employingrectangular geometry. FIG. 10 illustrates inlet contraction section1000, inlet minimum area 1002, subsonic diffuser section 1004,combustion chamber 1005, nozzle contraction section 1006, variablenozzle throat at the joining point of the contraction section 1006 andthe expansion section 1007. Step 1003 and the expansion section 1008 areillustrated in FIG. 10. All components of the dual-mode combustor 1000can vary dimensionally. In general the various components in FIG. 10 arerectangularly shaped. In this example, the nozzle throat would berectangular and would be adjustable.

FIG. 11 is a quarter-sectional diagrammatic view 1100 of the dual modecombustor 899 in the scramjet mode for flight Mach number 8 illustratinga radial step 1121A, a hinged diffuser section 1122, a hinged combustionsection 1123, a hinged contraction section 1124, a hinged nozzlethroat/arc section 1125 and a hinged expansion section 1126. FIG. 11Aprovides dimensional information 1100A relating to FIG. 11 including theradius of the engine at different stages thereof and the axial positionof different stages thereof. Reference numeral 1101 represents station 1(end of cylindrical inflow section), reference numeral 1102 representsthe beginning of cylindrical combustion chamber, reference numeral 1107represents station 7 (end of cylindrical combustion chamber), referencenumeral 1108 represents station 8 (nozzle throat), reference numeral1121 represents the cylindrical inflow chamber, reference numeral 1121Arepresents the hinge and aft facing step, reference numeral 1122represents the diffuser section, reference numerals 1122A, 1123A, 1127,1128 represent hinges, reference numeral 1123 represents the combustionchamber, reference numeral 1124 represents the contraction section,reference numeral 1125 represents the arc section, reference numeral1126 represents exhaust section, reference numeral 1129 represents thetermination of the exhaust section, reference numeral 1180 representsstation zero (air inlet from air inlet contraction device), referencenumeral 1180A represents the multi-mode fuel injectors, and referencenumeral 1181 indicates arrows of incoming air.

Also, hinges, H, indicate herein that the geometry of the dual-modecombustor may change around these points between component sectionsthereof to accommodate flight conditions. Reference numerals 1127 and1128 signify the interconnection of the arc section 1125 to thecontraction section and the expansion section, respectively. Inreviewing FIG. 11 tangency is maintained and required in all examplesbetween the arc sections and the contraction and expansion sections.This means that the hinges are the equivalent of sliding joints.Specifically, although joints 1127, 1128 are illustrateddiagrammatically as hinges, in fact these diagrammatic “hinges” arelimited in their movement such that tangency between the contractionsection and the arc section is maintained and the arc section may notbend back or extend such that a line coincident with the contractionsection would intersect with the arc section 1125. Similarly, the hingesillustrated in FIGS. 13-16, inclusive, may be considered as slidingjoints.

Still referring to FIG. 11, the hinges diagrammatically indicate thatthe geometry of the engine changes pursuant to the flight Mach numberconditions. Now referring to FIG. 12A, as a general rule thegeometric/nozzle throat 1202 and the inlet throat 1203 decrease withincreasing flight Mach number in ramjet mode. See FIG. 12. Similarly, asa general rule the geometric/nozzle throat 1205 and the inlet throat1203A decrease with increasing flight Mach number in scramjet mode. Theaxi-symmetric geometry used for the analysis consists of thefixed-length, hinged panels and cylindrical sections is shown in FIG.11. The fixed-length cylindrical inlet section diameter varies withflight Mach number to match the contraction ratio schedule given in FIG.12A with an allowance for fuel injection. A small radial step was placedat station 1 to facilitate flow separation. Generally the radial step isone-tenth the radius of the inlet cylinder. The cylindrical combustorsection is sized to accommodate ramjet combustion for the Mach 2.5flight condition. The nozzle throat is formed by a circular arc ofradius equal to one-half that of the inlet capture area. As the requiredthroat area varies with flight condition, the nozzle throat arc lengthvaries such that the contraction and expansion panels maintain tangency.The expansion panel trailing edge is maintained at a fixed radius,giving an exit area equal to twice the inlet capture area. Coordinatesfor the Mach 8 geometry shown in FIG. 11 are given in FIG. 11A. Ethylenefuel enters axially at station 1 (1101) through injectors 1180A asillustrated in FIG. 11.

FIG. 11B is a view of a receiving joint forming the nozzle throat.Reference numeral 1124B signifies a nozzle contraction section having areceiving joint 1125R. Reference numeral 1126B signifies a nozzleexpansion section having a receiving joint 1126R-receiving. Arc section1125B slidingly resides within joints/openings 1125R, 1126R such thatthe rotation of the nozzle contraction section 1124B and/or the rotationof the nozzle expansion section 1126R moves the nozzle throat 1108 whilemaintaining a tangential relationship between the sections 1124B, 1126Band the arc section 1125B.

FIG. 13A is a generalized quarter-sectional diagrammatic view 1300A ofthe flight Mach number 2.5 ramjet. FIG. 13B is a generalizedquarter-sectional diagrammatic view 1300B of the flight Mach number 3.0ramjet. FIG. 13C is a generalized quarter-sectional diagrammatic view1300C of the flight Mach number 4.0 ramjet.

All numerical values in FIGS. 13A-16C, inclusive, are in inches with theradius being indicated on the ordinate (“y”) axis and the axial lengthindicated on the abscissa (“x”) axis. Also, hinges, H, indicate hereinthat the geometry of the dual-mode combustor may change around thesepivot points between component sections thereof to accommodate flightconditions. Reference numerals H1 and H2 signify the interconnection ofthe arc section to the contraction section and the expansion section,respectively. In reviewing FIGS. 13A-16C, tangency is maintained andrequired in all examples between the arc sections and the contractionand expansion sections.

The reference numerals used in FIGS. 13A, 13B and 13C are set forthbelow. Reference numerals 1301I, 1311I, 1321I represent the respectiveinlet sections illustrated in FIGS. 13A, 13B and 13C, respectively.Reference numerals 1301A, 1311A, 1321A represent the arc sectionsillustrated in FIGS. 13A, 13B and 13C, respectively. Reference numerals1301N, 1311N, 1321N represent the variable nozzle throat sectionsillustrated in FIGS. 13A, 13B and 13C, respectively. A review of FIGS.13A, 13B and 13C, respectively, yields the conclusion that the inletdiametrical section, which is cylindrical, is decreasing in diameter asthe flight Mach number is increasing from 2.5 to 4.0 in the ramjet modewhile the nozzle throat radius is decreasing with increased flight Machnumber. Tangency is maintained in all examples of FIGS. 13A, 13B and 13Cbetween the arc sections and the contraction and expansion sections.

FIG. 14A is a generalized quarter-sectional diagrammatic view 1400A ofthe flight Mach number 5.0 ramjet. FIG. 14B is a generalizedquarter-sectional diagrammatic view 1400B of the flight Mach number 5.0scramjet. Reference numerals 1401I, 1411I represent the respective inletsections illustrated in FIGS. 14A and 14B, respectively. Referencenumerals 1401A, 1411A represent the arc sections illustrated in FIGS.14A and 14B, respectively. Reference numerals 1401N, 1411N represent thevariable nozzle throat sections illustrated in FIGS. 14A and 14B,respectively. A review of FIGS. 14A and 14B, respectively, yields theconclusion that the inlet diametrical section, which is cylindrical, isslightly increasing in diameter as the engine is transitioning fromramjet flight Mach number 5 to scramjet flight Mach number 5 while thenozzle throat radius is substantially increasing while transitioningfrom ramjet flight Mach number 5 to scramjet flight Mach number 5.Tangency is maintained in all examples of FIGS. 14A and 14B between thearc sections and the contraction and expansion sections.

FIG. 15A is a generalized quarter-sectional diagrammatic view 1500A ofthe flight Mach number 6.0 ramjet. FIG. 15B is a generalizedquarter-sectional diagrammatic view 1500B of the flight Mach number 6.0scramjet. Reference numerals 1501I, 1511I represent the respective inletsections illustrated in FIGS. 15A and 15B, respectively. Referencenumerals 1501A, 1511A represent the arc sections illustrated in FIGS.15A and 15B, respectively. Reference numerals 1501N, 1511N represent thevariable nozzle throat sections illustrated in FIGS. 15A and 15B,respectively. A review of FIGS. 15A and 15B, respectively, yields theconclusion that the inlet diametrical section, which is cylindrical, isslightly increasing in diameter as the engine is transitioning fromramjet flight Mach number 6 to scramjet flight Mach number 6 while thenozzle throat radius is substantially increasing while transitioningfrom ramjet flight Mach number 6.0 to scramjet flight Mach number 6.0.Tangency is maintained in all examples of FIGS. 15A and 15B between thearc sections and the contraction and expansion sections.

FIG. 16A is a generalized quarter-sectional diagrammatic view 1600A ofthe flight Mach number 8.0 scramjet. FIG. 16B is a generalizedquarter-sectional diagrammatic view 1600B of the flight Mach number 10.0scramjet. FIG. 16C is a generalized quarter-sectional diagrammatic view1600C of the flight Mach number 12.0 scramjet. Reference numerals 1601I,1611I, 1621I represent the respective inlet sections illustrated inFIGS. 16A, 16B and 16C, respectively. Reference numerals 1601A, 1611A,1621A represent the arc sections illustrated in FIGS. 16A, 16B and 16C,respectively. Reference numerals 1601N, 1611N, 1621N represent thevariable nozzle throat sections illustrated in FIGS. 16A, 16B and 16C,respectively. A review of FIGS. 16A, 16B and 16C, respectively, yieldsthe conclusion that the inlet diametrical section, which is cylindrical,is slightly decreasing in diameter as the flight Mach number isincreasing from 8.0 to 10.0 in the scramjet mode while the nozzle throatradius is moderately decreasing with increased flight Mach number.Tangency is maintained in all examples of FIGS. 16A, 16B and 16C betweenthe contraction and expansion sections.

Contours of static pressure ratio for flight Mach numbers 5, 8 and 12 inthe scramjet mode flight conditions appear in FIGS. 17A, 17B, and 17C.FIG. 17A is an illustration of the pressure contours 1700A within theengine for the flight Mach number 5.0 scramjet. FIG. 17B is anillustration of the pressure contours 1700B within the engine for theflight Mach number 8.0 scramjet. FIG. 17C is an illustration of thepressure contours 1700C within the engine for the flight Mach number12.0 scramjet.

Referring to FIG. 17A, pressure ratio contours, P/Pinlet, for the flightMach number 5.0 scramjet are illustrated and pressure ratio, P/Pinlet,1701, has a magnitude of about 1.04 and is located generally in therecirculation zone, forward portion of the combustion chamber. Referencenumeral 1731 is a stagnation streamline. When viewing FIG. 17A,everything leftwardly of stagnation streamline 1731 is in therecirculation zone. Reference numeral 1731A represents a free jetstreamline.

Referring to FIG. 17B, pressure ratio, P/Pinlet, 1711, for the flightMach number 8.0 scramjet, pressure ratio has a magnitude of about 1.32and is located generally in the recirculation zone of the forwardportion of the combustion chamber. When viewing FIG. 17B, everythingleftwardly of stagnation streamline 1732 is in the recirculation zone.Reference numeral 1732A represents a free jet streamline.

Referring to FIG. 17C, pressure ratio, P/Pinlet, 1712, for the flightMach number 12.0 has a magnitude of about 1.18 and is located generallyin the recirculation zone of the forward portion of the combustionchamber. When viewing FIG. 17C, everything leftwardly of stagnationstreamline 1733 is in the recirculation zone. Reference numeral 1733Arepresents a free jet streamline.

Reviewing FIG. 17, the recirculation zone pressure ratios increase fromscramjet flight Mach number 5 to 8 and then decrease from between flightMach number 8 to 12.

Contours of Mach number for flight Mach numbers 5, 8 and 12 in thescramjet mode flight conditions appear in FIGS. 18A, 18B, and 18C. Allthree cases for flight Mach numbers 5, 8 and 12 exhibit periodic wavestructure in the free-jet, and an overall increase in cross-sectionalarea due to combustion as the jet traverses the combustion chamber. Inall cases the free jet rejoins the nozzle throat contour and expands tothe exit area. The free-jet drives a primary recirculation zone in thecombustion chamber, the center of which moves aft with increasing flightMach number. Streamlines in the combustion chamber define therecirculation zone. In the inlet section, and continuing in a conicalnon-influence region of the free-jet, supersonic combustion elevates thepressure to a level higher than that of the reference pressure at theinflow plane. The highest pressure occurs on the axis, followed by anexpansion initiated at the jet boundary. In the Mach 8 and 12 cases, therecirculation zone equilibrates to the pressure at the radial step andthe bounding streamline issues axially with little initial deflection.In the Mach 5 case, the recirculation zone equilibrates to a lowerpressure, causing an initial expansion of the free-jet at the step. Allcases show a subsequent divergence of streamlines required toaccommodate the continuing supersonic combustion process while matchingcombustion chamber pressure. This “entry interaction” initiates therepetitive streamline structure characteristic of an under-expanded jet.The severity of the entry interaction depends on the initial rate ofmixing and combustion in the free-jet, and its initial pressure withrespect to the recirculation zone. The wavelength and shock lossesassociated with the streamline structure depend on the entryinteraction. At the combustor exit, the Mach 5 case approaches a soniccondition, and its wave structure disappears. Streamlines in the Mach 8case appear to be approximately in phase with the throat geometry, andthe streamlines merge smoothly into the minimum area. The Mach 12 casehowever, exhibits an “exit interaction” as streamlines are forced toconverge, resulting in a strong shock wave on the axis. This interactioncould obviously be eliminated by reducing the wavelength of the shockstructure or moving the throat, but of most benefit from a propulsionstandpoint would be to eliminate the periodic streamline structurealtogether by mitigating the entry interaction.

FIG. 18A is an illustration 1800A of the Mach number contours within theengine for the flight Mach number 5.0 scramjet. FIG. 18B is anillustration 1800B of the Mach number contours within the engine for theflight Mach number 8.0 scramjet. FIG. 18C is an illustration 1800C ofthe Mach number contours within the engine for the flight Mach number12.0 scramjet. Referring to FIG. 18A, reference numeral 1801 indicates amagnitude of about Mach 0.0 located in the recirculation zone of theforward portion of the combustion chamber.

Referring to FIG. 18B, reference numeral 1810 indicates a magnitude ofabout Mach 0.0 located in the recirculation zone in the middle of thecombustion chamber. Referring to FIG. 18C, reference numeral 1821represents a magnitude of about Mach 0.0 located in the recirculationzone of the aft portion of the combustion chamber.

FIG. 19A is an illustration of the static temperature contours 1900Awithin the engine for the flight Mach number 5.0 scramjet. FIG. 19B isan illustration of the static temperature contours 1900B within theengine for the flight Mach number 8.0 scramjet. FIG. 19C is anillustration of the static temperature contours 1900C within the enginefor the flight Mach number 12.0 scramjet.

Referring to FIG. 19A, reference numeral 1901 indicates a temperature ofabout 3500° R and reference numeral 1903 indicates a temperature ofabout 5000° R. Referring to FIG. 19B, reference numeral 1906 indicates atemperature of about 6000° R. Referring to FIG. 19C, reference numeral1910 indicates a temperature of about 9000° R.

Temperature contours appear in FIGS. 19A, 19B and 19C. The effects ofcombustion are apparent in the individual shear layers. The Mach 5 caseshows a degree of stratification that persists into the nozzle throat.The recirculation zone equilibrates to greater than 90% of theethylene-air theoretical value in the Mach 8 and 12 cases, but issignificantly cooler in the Mach 5 case. This is likely due to thetwo-injector arrangement used in the Mach 5 case, and suggests that therecirculation zone temperature and combustor heat load depend on thefuel injection method, and could be reduced in future design iterations.Exit interaction in the Mach 12 case may also contribute to elevatedtemperature in the recirculation zone.

In order to make a quantitative assessment of the losses in the free-jetcombustion process, and their effect on net thrust, mass-averaged axialdistributions of pressure, temperature, and velocity were obtainedduring the analysis. The combustor friction coefficient thus representsthe momentum loss associated with the recirculation zone and shockstructure in the free-jet. The ideal net thrust per unit airflow isillustrated in FIG. 20.

FIG. 20 illustrates the ideal net thrust per unit airflow based on useof different computational methods/tools. FIG. 20 illustrates ideal netthrust per unit of airflow against flight Mach numbers for aconventional ramjet, thermally-choked ramjet and a scramjet. Referencenumeral 2001 represents the ideal net thrust for scramjet modeoperation. Reference numeral 2003 represents the ideal net thrust forthe thermally choked operation such as in Curran et al. Referencenumeral 2002 represents the ideal net thrust for the ramjet disclosedherein. FIG. 20 illustrates a comparison of a thermally choked ramjet tothe dual-mode ramjet disclosed herein. The subsonic combustion ramjetdisclosed herein is 6-8% more efficient than the thermally-choked or“dual-mode” ramjet as a consequence of lower combustion Mach number. Ofgreater significance than higher performance however, is thepracticality of fuel distribution and flame-holding in the conventionalram burner.

FIG. 21 illustrates the mass-averaged static pressure distributions 2100with the pressure at the nozzle throat station denoted by symbols(supersonic combustor exit) for various flight conditions, to with,scramjet flight Mach numbers 5, 8 and 12. Reference numeral 2101represents Mach 5 pressure ratio data, reference numeral 2102 representsMach 8 pressure ratio data, and reference numeral 2103 represents Mach12 pressure ratio data. Compression due to mixing and combustion in thecylindrical inlet section from station zero to 0.36 feet is evident, asis the subsequent expansion and periodic streamline structure. As thefree-jet traverses the combustion chamber, the mean pressure isgenerally above the inflow value, consistent with the elevatedrecirculation zone pressures. The Mach 5 pressure distribution shows adamped character as combustion drives the free-jet toward a soniccondition. Of interest is the phase shift and elevated amplitude of thelast peak in the Mach 12 case consistent with the exit interaction seenin the pressure contours. Note that the combustor exit pressure (at theminimum area) used for cycle analysis of the Mach 8 and 12 solutions issignificantly higher than the inflow, and would cause a discrepancy withcycle analysis assuming combustion at constant pressure.

FIG. 21A illustrates the mass-averaged axial velocity ratio (V/V inlet)distributions 2100A for various flight conditions, to with, scramjetflight Mach numbers 5, 8 and 12. Reference numeral 2111 represents Mach5 velocity ratio data, reference numeral 2112 represents Mach 8 velocityratio data, and reference numeral 2113 represents Mach 12 velocity ratiodata. A marked reduction in velocity occurs upstream of the throatstation for the Mach 8 and 12 cases, and is more gradual for the Mach 5case, consistent with the pressure distributions. The loss coefficientsused to match the combustor exit velocities are listed in the FIG. 21A.Shock and viscous losses are represented in these values, and anestimate of their relative contributions to the total is not determined.Shock losses arise from the entry and exit interactions discussed above,and may be reduced by better tailoring of the combustion process, andoptimization of the combustion chamber geometry. The viscous loss arisesfrom the momentum required to drive the recirculating flow in thecombustion chamber, which presumably is a function of the combustionchamber volume and wetted area. These are determined by thecross-sectional area required at the minimum ramjet Mach number,subsonic diffuser length requirements, and the free jet length requiredfor supersonic mixing and combustion.

FIG. 21B illustrates the mass-averaged temperature distributions 2100Bfor scramjet mode flight Mach numbers 5, 8 and 12. Reference numeral2121 indicate Mach 5 temperature data as a function axial position,reference numeral 2122 indicate Mach 8 temperature data as a functionaxial position, and reference numeral 2123 represents Mach 12temperature data as a function of axial position. Temperatures increasewith increasing Mach flight numbers.

FIG. 23A illustrates the ethylene mass fraction 2300A for flight Machnumbers 5, 8 and 12 versus axial position. Reference numeral 2305signifies the flight Mach number 5, reference numeral 2306 signifies theflight Mach number 8, and reference numeral 2307 signifies the flightMach number 12.

Calculations at various nozzle throat areas were performed in order toevaluate the effect on recirculation zone pressure, entry and exitinteractions, and performance at the flight Mach number 8 as illustratedin FIGS. 11 and 16A. FIGS. 22A, 22B, 22C and 22D illustrate staticpressure contours for throat areas equal to 110%, 100%, 90% and 80% ofthe design value. FIG. 17B and FIG. 22B are identical but different datais presented and discussed in connection with each drawing figure.

FIG. 22A illustrates the static pressure ratio 2200A for scramjet modeflight Mach number 8 with the variable nozzle throat positioned at 110%of the design operating point. Reference numeral 2201 indicates astagnation streamline and reference numeral 2202 indicates the pressureratio of 0.95 located in recirculation zone of the combustion chamber(110% nozzle throat ratio). When viewing FIG. 22A, everything to theleft of stagnation streamline 2201 is in the recirculation zone.Reference numeral 2221T is the nozzle throat location (110% nozzlethroat area ratio).

FIG. 22B illustrates the static pressure ratio 2200B for scramjet modeflight Mach number 8 with the variable nozzle throat positioned at 100%of the design operating point. Reference numeral 2203 represents astagnation streamline and reference numeral 2204 indicates a pressureratio of 1.32 located in the recirculation zone of combustion chamber(100% nozzle throat ratio). When viewing FIG. 22B, everything to theleft of stagnation streamline 2203 is in the recirculation zone.Reference numeral 2223T is the nozzle throat location (100% nozzlethroat ratio).

FIG. 22C illustrates the static pressure ratio 2200C for scramjet modeflight Mach number 8 with the variable nozzle throat positioned at 90%of the design operating point. Reference numeral 2205 represents astagnation streamline and reference numeral 2206 is the pressure ratioof 1.60 located in recirculation zone of combustion chamber (90% nozzlethroat ratio). When viewing FIG. 22C, everything to the left and abovethe stagnation streamline 2205 is in the recirculation zone. Referencenumeral 2225T is the nozzle throat location (90% nozzle throat ratio).

FIG. 22D illustrates the static pressure ratio 2200D for scramjet modeflight Mach number 8 with the variable nozzle throat positioned at 80%of the design operating point. Reference numeral 2207 represents thestagnation streamline and reference numeral 2208 represents the pressureratio of 1.87 located in recirculation zone of combustion chamber (80%nozzle throat ratio). When viewing FIG. 22D, everything to the left andabove stagnation streamline 2207 is in the recirculation zone. Referencenumeral 2227T is the nozzle throat location (80% nozzle throat ratio).

As throat area is reduced, combustion chamber pressure increases, andthe period of the streamline structure decreases. As expected,combustion in the inlet section, and a short distance downstream is notaffected. Beyond this however, increased pressure increases the rate ofcombustion, reinforcing the tendency toward shorter wavelengths.Reference numerals 2201, 2203, 2205 and 2207 represent the streamlinesand streamline 2207 (variable nozzle throat at 80% of design value) hasa shorter wavelength than streamline 2201 (variable nozzle throat at110%) or streamline 2203 (variable nozzle throat at 100%). Further, thepressure increase in the combustion chambers is viewed in FIGS. 22A,22B, 22C and 22D as the variable nozzle's area is reduced. Referringback now to FIGS. 22A-D, it is evident that the free-jet entryconditions range from under-expanded at 110% throat area toover-expanded at 80%, but the streamline structure is never eliminateddue to the rapidity of combustion and divergence of streamlines in theinlet region. The severity of the exit interaction depends onsynchronization of the streamline structure with the throat geometry.The streamline 2203 associated with the variable nozzle throat at 100%of the design case appears to be in phase and exhibits almost no exitinteraction with the nozzle throat. Reference numeral 2223T representsthe variable nozzle throat for the 100% example. Reference numerals2221T, 2225T and 2207T represent the throats in the examples where thevariable nozzle throat is 110%, 90% and 80%, respectively. Interferencewith the throat is greatest for the 80 and 110% cases which show thestrongest interactions.

FIG. 23 illustrates the effect of nozzle throat area variation 2300 forscramjet mode flight Mach number 8 on the rate of ethylene fueldepletion. Reference numeral 2301 signifies the effect of throat areavariation on ethylene mass fraction (110% nozzle throat ratio),reference numeral 2302 signifies the effect of throat area variation onethylene mass fraction (100% nozzle throat ratio), reference numeral2303 signifies the effect of throat area variation on ethylene massfraction (90% nozzle throat ratio), and reference numeral 2304 signifiesthe effect of throat area variation on ethylene mass fraction (80%nozzle throat ratio).

FIG. 24 illustrates the effect of nozzle throat area variation onmass-averaged static pressure distribution 2400 for scramjet mode flightMach number 8. Reference numeral 2401 signifies the effect of throatarea variation on mass averaged static pressure distribution for theflight Mach number 8 (110% nozzle throat ratio), reference numeral 2402signifies the effect of throat area variation on mass averaged staticpressure distribution for the flight Mach number 8 (100% nozzle throatratio), reference numeral 2403 signifies the effect of throat areavariation on mass averaged static pressure distribution for the flightMach number 8 (90% nozzle throat ratio), and reference numeral 2404signifies the effect of throat area variation on mass averaged staticpressure distribution for the flight Mach number 8 (80% nozzle throatratio).

Mass-averaged pressure distributions for scramjet mode flight Machnumber 8 illustrated in FIG. 24 also show that as throat area isreduced, the initial pressure rise increases, the period of thestreamline structure decreases, and the mean is approximately equal tothe recirculation zone pressure. Peak-to-peak amplitude is roughly thesame for all cases. Note that for the 100% case, the waveform mergessmoothly with the nozzle expansion. The designation A8 in FIG. 24 refersto FIG. 11, station 8, reference numeral 1108. The 110% case shows aslight slope discontinuity just prior to the throat station and the 80and 90% cases show out-of-phase features at the throat, consistent withthe interactions seen in the pressure contours.

The effect of throat area variation was to change the combustion chamberpressure and the period of the streamline structure withoutsignificantly altering its amplitude. The amplitude of the primarystreamline structure is, therefore, most likely dependent on the initialrate of combustion. The exit interaction was affected by the phasing ofthe shock structure and was nearly eliminated in the 100% throat areacase. The adiabatic wall temperature and the gas temperature in therecirculation zone were not significantly affected by throat areavariation.

FIG. 25 illustrates 2500 the ideal net thrust per unit of airflowplotted against combustor exit pressure ratio and nozzle throat areavariation for scramjet mode flight Mach number 8. The ideal net thrustper unit airflow for the example of scramjet flight Mach number 8 forvariable nozzle throat opening ratios (80%, 90%, 100% and 110%) isplotted 2501 versus the mass-averaged combustor exit pressure ratio inFIG. 25. Reference numeral 2502 represents the ideal net thrust per unitairflow with a Cf of 0.0025. Friction loss coefficients required tomatch the exit velocities are also listed with the throat area for eachpoint. The 90% variable nozzle throat case exhibits the least momentumloss, the 110% case the greatest, and despite the entry and exitinteractions seen in pressure contours for the 80% case, its losscoefficient is slightly less than the 100% case which showed littleinteraction. This relative insensitivity and lack of correlation of losscoefficient to throat area is not unexpected however, since theamplitude of the basic streamline structure, and presumably the viscousloss component were not significantly affected. Cycle analysis resultsat the corresponding pressure ratios and with nominal momentum loss arealso plotted for reference and to show the basic sensitivity of scramjetnet thrust to combustor pressure ratio.

REFERENCE NUMERALS

Reference numerals 10-86 pertain to the prior art.

-   -   10—aircraft    -   12—ramjet combustion engine    -   14—inlet scoop    -   16—exhaust outlet    -   17, 18, 19—walls    -   20—fourth wall    -   21—converging inlet cowl passage    -   22—diverging supersonic combustion section    -   24—substantially uniform cross section subsonic combustion        section    -   26—exit nozzle    -   27—pilot zone recesses    -   28—fuel pump    -   30—fuel control system    -   32—plurality of nozzles    -   34—fuel control system    -   36—plurality of nozzles    -   40—central body    -   42—elongated inlet spike    -   43—flameholders    -   44—exhaust plug    -   46—annular member    -   47, 48—struts    -   49—fuel pump    -   50—subsonic combustion chamber    -   51—fuel control    -   52—nozzles 52 in the struts 47    -   55—nozzles supplied from fuel ducts    -   56—fuel ducts    -   58—recesses    -   60—supersonic combustion chamber    -   61—fuel control    -   62, 64—nozzles    -   65—ducts    -   70, 72—pumps    -   74, 75—nozzles    -   76, 78—fuel control system    -   80—supersonic combustion chamber    -   82—subsonic chamber    -   86—recess pilot zones    -   600—cross-sectional view of a prior art dual mode supersonic        ramjet engine operating in the scramjet mode    -   601—fuel injection nozzle    -   602—inlet contraction section    -   603—diverging supersonic combustion section    -   604—exit nozzle    -   605—fuel-air mixture    -   606, 606A, 880—incoming air being compressed    -   607, 607A, 881—exiting combustion gases    -   608—interior wall of engine    -   700—cross-sectional view of a prior art dual mode supersonic        ramjet engine operating in the thermally-choked ramjet mode    -   701—shock train to subsonic ramjet mode    -   702—beginning of shock train to subsonic ramjet mode    -   703—fuel injector    -   704—fuel injector

Reference numerals 800 and above pertain to the disclosed and claimedinvention.

-   -   800—perspective view of dual-mode combustor operating in the        ramjet mode    -   800A—cross-sectional schematic view of the dual-mode combustor        operating in the ramjet mode    -   800B—quarter sectional schematic view of the dual-mode combustor        operating in the ramjet mode    -   800C—enlarged portion of FIG. 8A illustrating the radial step        and the multimode fuel injector    -   801—inlet contraction section    -   802—inlet minimum area, variable diameter inlet cylindrical        passageway/section    -   803—radial step    -   804—subsonic diffuser section    -   805—combustion chamber    -   806—nozzle contraction section    -   807—variable nozzle throat at the joining point of the        contraction section 806 and the expansion section    -   808 in the ramjet mode or the scramjet mode    -   808—nozzle expansion section    -   810—ramjet mode flame holder    -   812—beginning of radial step 803    -   812A—end of radial step 803    -   830—terminal shock waves, position controlled by algorithm        governing nozzle throat position    -   841—supersonic compression    -   842—arrow indicating fuel injected    -   842I—multimode fuel injector    -   844—subsonic diffusion and fuel mixing    -   845—subsonic combustion    -   845A—supersonic combustion    -   846—contraction to choked throat    -   847, 847A—expansion    -   850—aperture in flame holder 810 for the passage of the free jet    -   872—heat release    -   899—dual-mode combustor    -   900—perspective view of dual-mode combustor operating in the        scramjet mode    -   900A—cross-sectional schematic view of the dual-mode combustor        operating in the scramjet mode    -   900B—quarter sectional schematic view of the dual-mode combustor        operating in the scramjet mode    -   900C—cross-sectional perspective view of the diffuser        illustrating the array of flame holders 810 and a central        aperture 850 within the array of flame holders 810    -   943—free-jet in the scramjet mode    -   943A—supersonic free jet boundary wherein the pressure is        approximately equal with that of the recirculation zone    -   944—recirculation zone    -   972—heat release    -   1000—perspective view of a dual-mode combustor using different        geometry    -   1001—inlet contraction section    -   1002—inlet minimum area    -   1003—step    -   1004—subsonic diffuser section    -   1005—combustion chamber    -   1006—nozzle contraction section    -   1007—variable nozzle throat at the joining point of the        contraction section 1006 and the expansion section 1008    -   1008—expansion section    -   1100—quarter sectional view of the dual-mode combustor in the        scramjet mode for flight Mach number 8    -   1100A—dimensional information for the quarter sectional view of        the dual-mode combustor in the scramjet mode for flight Mach        number 8    -   1100B—view of receiving joint forming the nozzle throat    -   1101—station 1, end of cylindrical inflow section    -   1102—station 2, beginning of cylindrical combustion chamber    -   1107—station 7, end of cylindrical combustion chamber    -   1108—station 8, nozzle throat    -   1121—cylindrical inflow chamber    -   1121A—hinge and aft facing step    -   1122—diffuser section    -   1122A, 1123A, 1127, 1128—hinge, sliding joint    -   1123—combustion chamber    -   1124—contraction section    -   1125—arc section    -   1126—expansion section    -   1124B—nozzle contraction section    -   1126B—nozzle expansion section    -   1126R—receiving joint    -   1125B—arc section    -   1125R—receiving joint    -   1129—termination of expansion section    -   1180—station zero, station i, air inlet from air inlet        contraction device    -   1180A—multi-mode fuel injectors    -   1181—arrows representing incoming air    -   1200—illustration of flight Mach number versus thermal throat        for prior art device, geometric/nozzle throat for dual-mode        combustor of present invention in ramjet mode and in scramjet        mode as a ratio of inlet capture area, and inlet throat in        ramjet mode and scramjet mode as a ratio of inlet capture area    -   1200A—table of flight Mach numbers versus inlet contraction        ratios, Ac/Ai    -   1200B—variable nozzle throat position schematic    -   1201—thermal throat of prior art device    -   1202—geometric/nozzle throat expressed as a ratio of nozzle        throat area to inlet capture area in ramjet mode    -   1203—dual mode combustor, inlet throat in ramjet mode    -   1203A—dual mode combustor, inlet throat in scramjet mode    -   1204—discontinuity/jump of variable nozzle throat position        between the ramjet mode 1202 and the scramjet mode 1205    -   1205—geometric/nozzle throat expressed as a ratio of nozzle        throat area to inlet capture area in scramjet mode    -   1206—desired ramjet nozzle throat position as a function of        flight Mach number for the ramjet mode    -   1207, 1209—switch    -   1208—desired ramjet nozzle throat position as a function of        flight Mach number for the scramjet mode    -   1210—controller operating on the difference of desired position        of the nozzle throat minus the actual position of the nozzle        throat    -   1211—output of controller    -   1212—nozzle throat positioner    -   1213—position signal    -   1214—variable geometric nozzle throat    -   1215, 1217—interconnecting signal transmission lines    -   1216—nozzle throat position sensor    -   1218—actual nozzle throat position as a function of flight Mach        number    -   1230—inlet contraction ratio    -   1231—combustion process    -   1300A—quarter-sectional schematic profile of the dual-mode        combustor in the ramjet mode, flight Mach number 2.5    -   1300B—quarter-sectional schematic profile of the dual-mode        combustor in the ramjet mode, flight Mach number 3    -   1300C—quarter-sectional schematic profile of the dual-mode        combustor in the ramjet mode, flight Mach number 4    -   1301A, 1311A, 1321A—arc section    -   1301C, 1311C, 1321C—combustion chamber    -   1301D, 1311D, 1321D—diffuser section    -   1301E, 1311E, 1321E—expansion section    -   1301I, 1311I, 1321I—inlet section    -   1301N, 1311N, 1321N—variable nozzle throat section    -   1301X, 1311X, 1321X—contraction section    -   1400A—quarter-sectional schematic profile of the dual-mode        combustor in the ramjet mode, flight Mach number 5    -   1400B—quarter-sectional schematic profile of the dual-mode        combustor in the scramjet mode, flight Mach number 5    -   1401A, 1411A—arc section    -   1401C, 1411C—combustion chamber    -   1401D, 1411D—diffuser section    -   1401E, 1411E—expansion section    -   14011, 1411I—inlet section    -   1401N, 1411N—variable nozzle throat section    -   1401X, 1411X—contraction section    -   1500A—quarter-sectional schematic profile of the dual-mode        combustor in the ramjet mode, flight Mach number 6    -   1500B—quarter-sectional schematic profile of the dual-mode        combustor in the scramjet mode, flight Mach number 6    -   1501A, 1511A—arc section    -   1501C, 1511C—combustion chamber    -   1501D, 1511D—diffuser section    -   1501E, 1511E—expansion section    -   1501I, 1511I—inlet section    -   1501N, 1511N—variable nozzle throat section    -   1501X, 1511X—contraction section    -   1600A—quarter-sectional schematic profile of the dual-mode        combustor in the scramjet mode, flight Mach number 8    -   1600B—quarter-sectional schematic profile of the dual-mode        combustor in the scramjet mode, flight Mach number 10    -   1600C—quarter-sectional schematic profile of the dual-mode        combustor in the scramjet mode, flight Mach number 12    -   1601A, 1611A, 1621A—arc section    -   1601C, 1611C, 1621C—combustion chamber    -   1601D, 1611D, 1621D—diffuser section    -   1601E, 1611E, 1621E—expansion section    -   1601I, 1611I, 1621I—inlet section    -   1601N, 1611N, 1621N—variable nozzle throat section    -   1601X, 1611X, 1621X—contraction section    -   1700A—pressure ratio, P/Pinlet, for the flight Mach number 5.0    -   1700B—pressure ratio, P/Pinlet, for the flight Mach number 8.0    -   1700C—pressure ratio, P/Pinlet, for the flight Mach number 12.0    -   1701—pressure ratio, P/Pinlet, about 1.04 located generally in        the forward portion of the combustion chamber    -   1711—pressure ratio, P/Pinlet, about 1.32 located generally in        the forward portion of the combustion chamber    -   1721—pressure ratio, P/Pinlet, about 1.18 located generally in        the forward portion of the combustion chamber    -   1731, 1732, 1733—stagnation streamline    -   1731A, 1732A, 1733A—free-jet streamline    -   1800A—Mach number contours for the flight Mach number 5.0    -   1800B—Mach number contours for the flight Mach number 8.0    -   1800C—Mach number contours for the flight Mach number 12.0    -   1801—about Mach 0.0, located in the recirculation zone of the        forward portion of the combustion chamber    -   1810—about Mach 0.0, located in the recirculation zone in the        middle of the combustion chamber    -   1821—about Mach 0.0, located in the recirculation zone of the        aft portion of the combustion chamber    -   1900A—static temperature contours for the flight Mach number 5.0    -   1900B—static temperature contours for the flight Mach number 8.0    -   1900C—static temperature contours for the flight Mach number        12.0    -   1901—3500° R    -   1903—5000° R    -   1906—6000° R    -   1910—9000° R    -   2000—ideal net thrust per unit airflow over various flight Mach        numbers    -   2001—scramjet mode net thrust    -   2002—conventional, prior art, net thrust in the ramjet mode    -   2003—Curran (prior art) ramjet mode net thrust    -   2100—mass averaged pressure distributions for scramjet flight        mach numbers 5, 8 and 12    -   2100A—mass averaged axial velocity distributions for scramjet        flight mach numbers 5, 8 and 12    -   2100B—mass averaged temperature distributions for scramjet        flight mach numbers 5, 8 and 12    -   2101—Mach 5 pressure ratio data as a function of axial position    -   2102—Mach 8 pressure ratio data as a function of axial position    -   2103—Mach 12 pressure ratio data as a function of axial position    -   2111—Mach 5 axial velocity ratio data as a function of axial        position    -   2112—Mach 8 velocity ratio data as a function of axial position    -   2113—Mach 12 velocity ratio data as a function of axial position    -   2121—Mach 5 temperature data as a function of axial position    -   2122—Mach 8 temperature data as a function of axial position    -   2123—Mach 12 temperature data as a function of axial position    -   2200A—static pressure plot for variable area nozzle throat        position at 110% of design point for the flight Mach number 8    -   2200B—static pressure plot for variable area nozzle throat        position at 100% of design point for the flight Mach number 8    -   2200C—static pressure plot for variable area nozzle throat        position at 90% of design point for the flight Mach number 8    -   2200D—static pressure plot for variable area nozzle throat        position at 80% of design point for the flight Mach number 8    -   2201—stagnation streamline    -   2202—pressure ratio of 0.95 located in recirculation zone of        combustion chamber (110% nozzle throat ratio)    -   2203—stagnation streamline line    -   2204—pressure ratio of 1.32 located in recirculation zone of        combustion chamber (100% nozzle throat ratio)    -   2205—stagnation streamline line    -   2206—pressure ration of 1.60 located in recirculation zone of        combustion chamber (90% nozzle throat ratio)    -   2207—stagnation streamline line    -   2208—pressure ratio of 1.87 located in recirculation zone of        combustion chamber (80% nozzle throat ratio)    -   2221T—nozzle throat location (110% nozzle throat ratio)    -   2223T—nozzle throat location (100% nozzle throat ratio)    -   2225T—nozzle throat location (90% nozzle throat ratio)    -   2227T—nozzle throat location (80% nozzle throat ratio)    -   2300—effect of throat area variation on ethylene mass fraction        for the flight Mach number 8    -   2300A—ethylene mass fraction for scramjet mode flight Mach        numbers 5, 8 and 12 versus axial position    -   2301—effect of throat area variation on ethylene mass fraction        (110% nozzle throat ratio)    -   2302—effect of throat area variation on ethylene mass fraction        (100% nozzle throat ratio)    -   2303—effect of throat area variation on ethylene mass fraction        (90% nozzle throat ratio)    -   2304—effect of throat area variation on ethylene mass fraction        (80% nozzle throat ratio)    -   2305—Mach flight number 5 axial position and ethylene mass        fraction    -   2306—Mach flight number 8 axial position and ethylene mass        fraction    -   2307—Mach flight number 12 axial position and ethylene mass        fraction    -   2400—effect of throat area variation on mass averaged static        pressure distribution for the flight Mach number 8    -   2401—effect of throat area variation on mass averaged static        pressure distribution for the flight Mach number 8 (110% nozzle        throat ratio)    -   2402—effect of throat area variation on mass averaged static        pressure distribution for the flight Mach number 8 (100% nozzle        throat ratio)    -   2403—effect of throat area variation on mass averaged static        pressure distribution for the flight Mach number 8 (90% nozzle        throat ratio)    -   2404—effect of throat area variation on mass averaged static        pressure distribution for the flight Mach number 8 (80% nozzle        throat ratio)    -   2500—ideal net thrust per unit airflow as a function of nozzle        throat pressure ratio, Pnozzle/Pinlet    -   2501—net thrust per unit airflow for the current free-jet        disclosed herein    -   2502—net thrust per unit airflow with a Cf of 0.0025.    -   A=Cross-sectional area    -   Cf=Friction coefficient    -   D=Nozzle throat diameter ramjet mode    -   D1=Nozzle throat diameter scramjet mode    -   H=Hinge/sliding joint    -   H1=First Arc Hinge/sliding joint    -   H2=Second Arc Hinge/sliding joint    -   M=Mach number    -   P=Pressure    -   r=Radial distance    -   x=Axial distance    -   Z=Altitude

SUBSCRIPTS

-   -   0=Freestream    -   1=Cylindrical inflow section exit station    -   2=Combustion chamber inlet station    -   7=Combustion chamber exit station    -   8=Nozzle throat station    -   C=Inlet capture area    -   i=Inflow station    -   min=Minimum    -   T=Total

Those skilled in the art will readily recognize that the invention hasbeen set forth by way of example only and that changes may be made tothe examples without departing from the spirit and the scope of theclaims which follow herein below.

The invention claimed is:
 1. A process for operating a dual-modecombustor, comprising the steps of: using an inlet contraction chamberto compress combustion air into a combustion air passageway; injectingfuel from said combustion air passageway into said combustion air insaid combustion air passageway creating a fuel-air mixture; feeding saidfuel-air mixture from said combustion air passageway into a diffusionsection; igniting said fuel-air mixture in said diffusion section andcombusting said fuel-air mixture in said diffusion section and in acombustion chamber, where a portion of the combusting fuel-air mixtureengages a flame holder affixed to the combustion chamber to stabilizethe combusting fuel-air mixture; evacuating said combusted fuel-airmixture from said combustion chamber and into a variable area nozzlethroat; and, modulating and positioning said variable area nozzle throataccording to an algorithm creating and controlling the position of aterminal shock within said diffusion section, said algorithm being afunction of flight Mach number; and further comprising the step of notengaging said flame holder with said portion of the combusting fuel-airmixture in a scramjet mode, and wherein said step of modulating andpositioning a variable area nozzle throat according to an algorithmcreating and controlling said position of said terminal shock withinsaid diffusion section, includes transitioning, using said algorithm,said dual-mode combustor from a ramjet mode to said scramjet mode byopening said variable area nozzle throat at a specified flight Machnumber.
 2. A process for operating a dual-mode combustor as claimed inclaim 1 wherein said algorithm includes a discontinuity where there aretwo values for a specified flight Mach number.
 3. A process foroperating a dual-mode combustor as claimed in claim 2 wherein saidfuel-air mixture and said combusted fuel-air mixture is separated into afree-jet from said combustion air passageway where a step is located tosaid variable area nozzle throat.
 4. A process for operating a dual-modecombustor as claimed in claim 3 wherein said free-jet does not engagesaid subsonic diffuser, said combustion chamber or a contractionsection.
 5. A process for operating a dual-mode combustor as claimed inclaim 4 wherein said algorithm includes said variable area nozzleposition as a ratio A/Ac of the actual nozzle throat area, A, to theinlet capture area, Ac, of said contraction section, said variable areanozzle throat position varies from a ratio of about 0.8=A/Ac at aboutflight Mach number 2.5 in the ramjet mode to a ratio of about 0.18=A/Acat about flight Mach number 5.0 in the ramjet mode, said nozzle throatposition varies rapidly from a ratio of about 0.18=A/Ac at about flightMach number 5.0 in the ramjet mode to a ratio of about 0.41=A/Ac atabout flight Mach number 5.0 transitioning to the scramjet mode,thereafter, said nozzle throat position varies from about 0.41=A/Ac atflight Mach number 5.0 in the scramjet mode to a ratio of about0.15=A/Ac at about flight Mach number 12 in the scramjet mode.
 6. Aprocess for operating a dual-mode combustor as claimed in claim 4,wherein said algorithm includes reducing said variable area nozzlethroat from an initial value to a first value from about 2.5-5.0 flightMach number in the ramjet mode, enlarging said open area of saidvariable area nozzle throat at about flight Mach number 5.0 from saidfirst value to a second value, said second value representing a largerarea than said first value, transitioning from said ramjet mode to saidscramjet mode, and then, reducing said open area of said variable areanozzle throat from said second value representing the larger area to afinal value from about 5.0-12.0 flight Mach number in the scramjet mode.